FIG. 1A is a simplified partial cutaway view of a prior art gas turbofan engine 1000 (sometimes referred to as the “engine” for brevity) having a rotational axis A. The engine 1000 includes an air intake 1100, a propulsive fan 1200, an intermediate-pressure compressor 1300, a high-pressure compressor 1400, a combustor 1500, a high-pressure turbine 1600, an intermediate-pressure turbine 1700, a low-pressure turbine 1800, and an exhaust nozzle 1900. The high-pressure compressor 1400 and the high-pressure turbine 1600 are connected via a shaft 2000 and rotate together about the rotational axis A. The intermediate-pressure compressor 1300 and the intermediate-pressure turbine 1700 are connected via a shaft 2100 and rotate together about the rotational axis A. The fan 1200 and the low-pressure turbine 1800 are connected via a shaft 2200 and rotate together about the rotational axis A. A fan nacelle 2400 generally surrounds the fan 1200 and defines the air intake 1100 and a bypass duct 2300. Fan outlet guide vanes 2500 secure the fan nacelle 2400 to the core engine casing.
In operation, the fan 1200 compresses air entering the air intake 1100 to produce a bypass air flow that passes through the bypass duct 2300 to provide propulsive thrust and a core air flow into the intermediate-pressure compressor 1300. The intermediate-pressure compressor 1300 compresses the air before delivering it to the high-pressure compressor 1400. The high-pressure compressor 1400 further compresses the air and exhausts the compressed air into the combustor 1500. The combustor 1500 mixes the compressed air with fuel and ignites the fuel/compressed air mixture. The resultant hot combustion products then expand through—and thereby drive—the high-, intermediate-, and low-pressure turbines 1600, 1700, and 1800 before being exhausted through the exhaust nozzle 1900 to provide additional propulsive thrust. The high-, intermediate-, and low-pressure turbines 1600, 1700, and 1800 respectively drive the high-pressure compressor 1400, the intermediate-pressure compressor 1300, and the fan 1200 via the respective shafts 2000, 2100, and 2200.
FIG. 1B shows a simplified cross-sectional view of the engine 1000 taken along a plane perpendicular to the axis A through a nozzle guide vane assembly between the high-pressure turbine 1600 and the intermediate-pressure turbine 1700. The engine 1000 includes a turbine section casing 10 including an outer cylindrical wall 11 and an inner cylindrical wall 12 radially inward of the outer wall 11 that together define a cooling air chamber 13 there between. An inner flow path boundary wall 16 and inner cylindrical wall 12 define the inner and outer boundaries of the working fluid flow path, respectively. The inner flow path boundary wall 16 defines a turbine chamber 14. Engine shafts, rotor discs and bearings may be contained within the turbine chamber 14 and require temperature control. Multiple conduits 20a-20i are circumferentially arranged around the turbine chamber 14 and about the axis A and extend radially inwardly from respective conduit inlets on the inner cylindrical wall 12 in fluid communication with the cooling fluid chamber (manifold) 13 to respective conduit outlets on the inner flow path boundary wall 16 in fluid communication with the turbine chamber 14. The conduits 20a-20i are positioned through nozzle guide vanes (not shown) of the nozzle guide vane assembly. The cooling fluid chamber 13 is fluidically connectable to a cooling fluid source (such as the compressor stage of the engine 1000) via inlet tubes 1, 2, and 3. It is conventional in the prior art that the number of conduits is divisible by the number of inlets, or in other words the number of conduits is an integer multiple of the number of inlets. The inlet tubes 1, 2, and 3 are circumferentially spaced about the axis A.
In operation, cooling air flows from the cooling air source through the inlet tubes 1, 2, and 3 and into the cooling air chamber 13. The cooling air then flows from the cooling air chamber 13 through the conduits 20a-20i into the turbine chamber 14. The cooling air may also cool the nozzle guide vanes as it flows through the conduits 20a-20i and then cools the components in the turbine chamber 14.
The circumferential positions of the conduits 20a-20i relative to the inlet tubes 1, 2, and 3 and the fact that the conduits 20a-20i are identical causes non-uniform flow distribution in the cooling air chamber 13 and through the conduits 20a-20i. In other words, the mass flow rates of the cooling air through the respective conduits 20a-20i are not substantially uniform. FIG. 1C is a graph showing the mass flow rate distribution (curve F) and the static pressure distribution (curve P) of the cooling air in the cooling air chamber 13 relative to the circumferential angle about the axis A. Generally, the mass flow rate increases and the static pressure decreases when based on increasing proximity to one of the inlet tubes 1, 2, and 3 is located, and vice-versa.
This non-uniform flow distribution can cause temperature variations within the turbine section casing 10 that can lead to components cracking or overheating. It can also cause non-uniform turbine section casing growth that could affect the spacing between the tips of the turbine blades and the inner wall 12 of the turbine section casing 10. Non-uniform flow distribution can also lead to insufficient cooling of certain nozzle guide vanes due to less cooling air flow through their respective conduits. This can result in nozzle guide vane burn through, i.e., hole formation in the nozzle guide vanes which disrupts flow and introduces cooler air into the working fluid and thus decreasing the efficiency of the engine. Non-uniform flow distribution can also lead to uneven heat pickup from the gas path to the cooling air as it passes through the nozzle guide vane, leading to an uneven temperature distribution within the turbine chamber 14.